1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to an air cooled turbine rotor blade in a gas turbine engine.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream. In engines of the future, it is even anticipated that third stage airfoils will also require cooling such as to prevent erosion and limit creep.
In an industrial gas turbine (IGT) engine, the turbine is designed to withstand the highest turbine inlet temperature that can be operated while allowing for the turbine to run constantly under these conditions for long periods of time. Airfoil cooling is performed so that an airfoil mass average sectional metal temperature does not exceed a certain temperature in order to improve airfoil creep capability for a turbine rotor blade. Creep is when the blade stretches in length due to the high radial stress loads produced from the blade rotating while exposed to the high temperatures. As the metal temperature increases, the metal becomes weaker and can become over-stressed. The gap spacing between the blade tips and the outer shroud must be kept to a minimum to control blade tip leakage. When a blade creep occurs, the gap can become negative such that excessive rubbing will occur.
Prior art airfoil cooling makes use of a triple pass (3-pass) serpentine flow cooling circuit that includes a forward flowing triple pass serpentine circuit 10 and an aft flowing serpentine circuit 20. The forward flowing triple pass serpentine circuit 10 includes a first leg 11, a second leg 12 and a third leg 13 that is connected to the leading edge impingement channel or cavity 15 through a row of metering and impingement holes. The showerhead arrangement of film cooling holes (three film holes) and two gill holes (one of the P/S and another of the S/S) discharge film cooling air from the spent impingement cooling air in the L/E channel 15. The forward flowing circuit 10 normally is designed in conjunction with leading edge backside impingement cooling plus a showerhead arrangement of film cooling holes with pressure side and suction side gill holes to provide cooling for the leading edge region of the blade.
The aft flowing serpentine flow circuit 20 is designed in conjunction with the airfoil trailing edge discharge cooling holes. This type of cooling flow circuit is called a dual triple pass serpentine “warm bridge” cooling design with three legs 21-23 and is shown in FIGS. 1 and 2. No film cooling holes are used along the middle section of the airfoil that discharges film cooling air from the serpentine flow cooling circuit. The “warm bridge” cooling circuit operates as follows. Cooling air flows into the forward flowing serpentine circuit 10 in a first leg 11 towards the blade tip, then turns into a second leg 12 and flows toward the root, and then flows into a third leg 13 toward the blade tip, where the third leg 13 is adjacent to the leading edge impingement cavity 15 so that cooling air is bled off through a row of metering and impingement holes to produce impingement cooling against the leading edge wall, in which the spent impingement cooling air then flows out through the showerhead film cooling holes. The aft end side of the blade is cooled with an aft flowing triple pass serpentine circuit 20 and flows through the three legs 21-23 in which the third leg 23 is located adjacent to the trailing edge region. The cooling air from the third leg 23 flows through trailing edge exit holes to cool the trailing edge region.
An alternative prior art cooling design to that of FIGS. 1 and 2 utilizes the dual triple pass serpentine flow circuits for a high operating gas temperature and is shown in FIGS. 3 and 4. The FIGS. 3 and 4 blade cooling circuit is called a “cold bridge” cooling design. In this “cold bridge” cooling circuit, the leading edge airfoil is cooled with a self-contained flow circuit 31. The airfoil mid-chord section is cooled with a triple pass serpentine flow circuit 32. The trailing edge region is cooled with a triple-pass forward flowing serpentine cooling circuit 33 that continues toward the mid-chord triple pass serpentine flow circuit 32. However, the aft flow circuit is flowing in a forward direction instead of the aftward direction as in the “warm bridge” design of FIGS. 1 and 2. Again, the aft flowing serpentine flow circuit is designed in conjunction with the airfoil trailing edge discharge cooling holes. FIG. 4 shows a flow diagram for this “cold bridge” cooling circuit which has two forward flowing triple pass serpentine flow circuits 32 and 33 plus a leading edge cooling air supply channel 31 separate from the triple pass serpentine flow circuits that is used for cooling the leading edge region and discharging the film cooling air through the showerhead holes.
In both of these prior art blade serpentine flow cooling circuits, the internal cavities are constructed with internal ribs that extend across the channels and connect the airfoil pressure side and suction side walls. In most cases, the internal cooling cavities are at a low aspect ratio which is subject to high rotational effect on the cooling side heat transfer coefficient. In addition, the low aspect ratio cavity yields a very low internal cooling side convective area ratio to the airfoil hot gas external surface.